Rocket engine with thrust direction modifying devices

ABSTRACT

1. A rocket engine comprising, in combination, a manifold for introducing propellants into a combustion chamber, a wall having a hemispherical deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound, a conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, apertures formed in the apex of said chamber for bleeding combustion gases therethrough, and means for deflecting a portion of said gases to vary the thrust vector of said engine in operation.

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"2542 XR 3963617 1 3,636,710 V W- Jan. 25, 1972 [54] ROCKET ENGINE WITH THRUST gar; ..60/35.6 ry en 60/35 6 DIRECTION MODIFYING DEVICES 3,145,530 8/1964 Sabey 60/35 6 [72] Inventors: Joseph J. Lovingham, Madison; Hartmann 1,446 10/1964 Parilla 60/35 6 J. Kircher, III, Sparta, both of NJ. 3,173,250 3/1965 Matzenauer.. 60/35 6 3,192,714 7 1965 H k .6 [731 Assignee Chemical Cmpmfim, 3 216 191 11/1965 M211 5; al. "63132.2 [22] Filed: Oct. 22, 1965 Primary Examiner-Samuel Feinberg PP N05 501,063 Att0rney-William R. Wright, Jr.

Related US. Application Data EXEMPLARY CLAXM [62] 5 2 3 1962 1. A rocket engine comprising, in combination, a manifold for introducing propellants into a combustion chamber, a wall having a hemispherical deflecting surface fixed to an inter- [52] US. Cl 660002235? mediate portion of Said manifold and extending aft 5] l t Cl d 1/14 therearound, a conical combustion chamber fixed to and com- "gs 35g 231 230 municating with the aft end of said manifold and having its 1 o are larger open end positioned adjacent but spaced from said sur- 56 R f d face to define an expansion exhaust nozzle and throat 1 e erences l e therewith, apertures formed in the apex of said chamber for UNITED STATES PATENTS bleeding combustion gases therethrough, and means for deflecting a portion of sad gases to vary the thrust vector of 2,434,652 1/1948 Hickman .60/35 6 aid engine in operation 2,500,117 3/1950 Chandler (SO/35.6 3,016,697 1/1962 Sternberg et al... 60/356 17 Chums, 27 Drawing Figur s 3,091,081 5/1963 Alperet a1 60 3 5 6 PATENTEI] mas-1972 3636710 sum 01! 17 PATENTEDJANZSIBYZ 3.836.710

sum near 11 PATENTED m2 5 m2 sum 03 or n ZOEHEZKMP ZEPmDm PATENTED JAN25 I372 3.636.710

' sum USUF 17 INVENTORS JOSE H J. LOVING/MM PATENTED M25 ISYZ 31536.7 1 0 SHEET 08 0F 17 ACORN DEFLECTE D FREE SURFACE ACORN coueusnou CHAMBER 36 MACH LINE-/" INVENTORS JOSfPl-l J. LOV/NG/MM HART/MAW J. K/PCl/[P E AGE PATENTEDJANZSIQYZ 3.636.710 sum near 17 4! THROAH ensue) mnonnnnor use) FREE SURFACE 60- ACTUAI'ING CYLINDER ACORN couaus'rlou mesa-36 mu uus 59' Piston cormscreo to ACORN INVENTORS JOSEPH J. l0W/V6l/AM BY HAFIMANNJK/KHERM PATENTEDJANZSISTZ 3,636,710

SHEET lOUF 17 THROAT (BASIC) 2 35 ACORN DEFLECTED 63-THROAT (THROTTLED) FREE SURFACE SHOCK 60 ACTUAI'ING cvuuoen ACORN COMBUSTION CHAMBER 36 MACH LlNE'\ PATENTEDJANZ'SIHYZ 3.636.710

SHEET mar 17 SHOCK FREE SURFACE ACORN comausnon CHAMBER-36 MACH LINE\ Fly 12 AGFNT PATENTEUJANZSIQYZ 3.636.710

SHEET I 1W 1? ACORN CHAMBER 57 CATALYST BED' 88 Fly. A9

INVENTORS' JOSEPH J. LOVl/Vl/AM HA QTMA NA/ J. A M CHEF H AGENT PATENTEU JAN25 I972 3.636.710

SHEET lBUF 17 [SUSTAIN PHASE POSITION BOOST PHASE POSIT'ON INVENTORS JOYFPIV J. ZOV/NG/MM HAATMA/VA/ J K/PCHEA E l AGE/V ROCKET ENGINE WITH TI-IRUST DIRECTION MODIFYING DEVICES This application is a division of Ser. No. 247,443, filed Dec. 18, 1962, and entitled Missile and Powerplant and now US. Pat. No. 3,482,404, granted Dec. 9, 1969.

This invention relates generally to reaction motors and particularly to an improved rocket powerplant or missile having an improved construction and a novel type of thrust chamber.

Rocket powerplants of various types are well known in the art and most of these are characterized by at least several of a considerable number of disadvantages inherent in each of the various types. Among these disadvantages are: the use of separate thrust chambers for multistage operation; a lowered payload or fuel capacity due to the volume of the powerplant occupied by the thrust chamber; an inability to vary the thrust of the powerplant or to effect thrust vector control; the use of complicated control and/or pressurizing systems resulting in added weight and airframe space requirements; an inability to efi'ect packaging of the powerplant; an inability to efi'ect a termination of thrust simply and accurately; and an excessive weight, length and/or cost without resulting in a higher reliability.

Accordingly the main object of the present invention is to provide an improved, packaged, rocket powerplant having an engine with a novel type of thrust chamber and which will obviate the above and other disadvantages characterizing known prior art structures.

An important object of the present invention is to provide an improved rocket powerplant in which the volume normally occupied by the combustion chamber of the engine may be utilized for propellant loading to effect an increase in mass ratio for the missile.

Another important object of the present invention is to provide an improved rocket powerplant having a rocket engine embodying a novel thrust chamber hereinafter designated as an acom" thrust chamber which provides: a control of thrust level; an optimum nozzle expansion; a simple, high response, thrust vector control; efficient packaging; and simple, accu rate thrust termination.

A further important object of the present invention is to provide an improved rocket powerplant and missile which; has a rocket engine which uses a single combustion chamber for both boost and sustain thrust; uses a single, low-temperature solid propellant gas generator for propellant feed, thrust vector control, roll control and Vernier velocity control; employs jettisonable booster tanks that are separated by a liquid fueled, shaped charge system that is inert until the engine ignites; and contains the payload, guidance, etc., within the tankage, and thus affords the advantages of lower weight, shorter length, lower cost, and higher reliability, over the conventional two stage approach.

A still further important object of the present invention is to provide an improved rocket powerplant of the type described which embodies a booster tank jettison system, a safety des'truct system, and which may be armed when desired by the insertion of initiators for the two systems and for the solid pressurization grain of the gas generator.

Another important object of the present invention is to provide an improved rocket powerplant having a novel engine of the type described wherein thrust vectoring during boost operation is achieved by swiveling the acorn thrust chamber about a universal joint by mechanical means or by means of combustion gases bled through the aft end of the chafinber and controlled by the oblique shock technique.

A further important object (if the present invention is to provide a fluid pressure balanced, universal joint pivot for the acorn thrust chamber of the engine so that only minimal forces are required to effect swiveling of the chamber during thrust vectoring.

A still further important object of the present invention is to provide novel means for efiecting thrust vectoring of a rocket engine having a fixed acorn thrust chamber.

Another important object of the present invention is to provide a two stage, packaged, liquid propellant, powerplant having a solid propellant pressurizing system embodying automatic valves operable upon termination of the first stage to shut off pressurizing gas flow to the first stage propellant tanks and admit it to the second stage tanks.

A further important object of the present invention is to provide a two stage, packaged, liquid propellant powerplant having a propellant flow control valve assembly initially operable by propellant pressure for first stage propellant flow, then operable by pressuring gases to second stage operation, and finally operable to shut down position by spring means.

A still further important object of the present invention is to provide a novel rocket engine having a thrust chamber of conical or acom" shape with an open base or forward end positioned adjacent a contoured deflecting surface with which the rim of the cone or acom defines an expansion nozzle for the combustion gases which burn in the acom.

Another important object of the present invention is to provide a missile in the form of an improved rocket powerplant which is the vehicle for a payload or warhead, and a guidance system, etc., mounted within the propellant tankage; employs jettisonable booster tanks; and is provided with a rocket engine having a single combustion chamber for both booster and sustain operational phases.

Other objects and advantages of the present invention will become apparent during the course of the following description.

In the drawings we have shown several embodiments of the invention. In these showings:

FIG. 1 is a schematic view of the novel rocket powerplant and missile comprising the present invention;

FIG. 1A is a similar fragmentary view showing the amplified thrust vector effect when the acorn thrust chamber is tilted;

FIG. 2 is a schematic view to a reduced scale showing the several stages of operation of the invention;

FIG. 3 is a schematic view of the preferred form of the novel rocket powerplant showing the propellant, pressurizing, and control connections;

FIG. 4 is a fragmentary, central, longitudinal sectional view to an enlarged scale of the preferred form of the acorn thrust chamber and nozzle and the controls therefor;

FIG. 5 is a fragmentary similar view thereof but taken at an angle turned with respect to FIG. 4;

FIG. 6 is a schematic, fragmentary, central longitudinal sectional view of a fixed acorn thrust chamber and nozzle as used with bipropellants;

FIG. 7 is a similar view showing the acorn thrust chamber provided with one form of mechanical means for effecting thrust vector control;

FIG. 8 is an aft end view of the vector control sectors of FIG. 7;

FIG. 9 is a view similar to FIG. 7 showing a modified fonn of vector control means;

FIG. 10 is a view similar to FIG. 6 but showing the acom" thrust chamber and nozzle swivelly mounted by means of a ball joint pivot and another modification of mechanical means for effecting vector control by tilting the acorn;

FIG. 11 is a view similar to FIG. 6 but showing the acorn thrust chamber mounted for axial movement by fluid power means to vary the throat area and hence the thrust of the powerplant;

FIG. 11A is a view similar to FIG. 6 but showing the acorn chamber swivelly mounted to mechanically vary the thrust vector, and for axial movement;

FIG. 12 is a view similar to FIG. 10 but showing the employment of aerodynamic actuation for thrust vector control, combustion gases being bled from a nozzle at the apex of the acorn rearwardly and subject to auxiliary moment producing control forces; FIG. 13 is a fragmentary sectional view to an enlarged scale of the acorn and its apex nozzle of FIG. 12 showing the means for producing the auxiliary control forces;

FIG. 14 is a transverse, sectional view thereof taken on the line 14-14 of FIG. 13;

FIG. is a fragmentary schematic view showing the axial nature of the flow through the apex nozzle in the absence of control fluid forces;

FIG. 16 is a similar view showing the deflection of the apex nozzle flow under the influence of control fluid;

FIG. 17 is a central, longitudinal sectional view to a decreased scale of another form of the acorn powerplant which uses a solid propellant;

FIG. 18 is a similar view of another form of the acorn powerplant which is a hybrid using both liquid and solid propellants;

FIG. 19 is a similar fragmentary view of another form of the acom powerplant which uses a monopropellant and a catalyst bed;

FIG. 20 is a transverse sectional detail view taken on the line 2020 of FIG. 1;

FIG. 21 is a similar view taken on the line 2121 of FIG. 1;

FIG. 22 is a similar view taken on the line 22-22 of FIG. 1;

FIG. 23 is a greatly enlarged view of the dash line encircled portion of FIG. 1;

FIG. 24 is an enlarged view of the flow control valve assembly portion of the controls shown in FIG. 4; and

FIG. 25 is a schematic view of the signaling circuits for effecting the proper sequential operational steps of the powerplant.

Referring to the drawings, the preferred form of the improved missile and powerplant which is designated as a whole by numeral 30, includes a sustain fuel tank SF, a sustain oxidi'zer tank SO, a gas generator chamber PG, and an abutting cylindrical chamber G containing guidance controls, etc., all being rigidly connected as an airframe designated as a whole by numeral 32.

A payload P which may be a warhead, instrumentation, and/or personnel capsule, guidance, etc., is releasably mounted on the forward end of the chamber G and enclosed by annular, jettisonable boost oxidizer and boost fuel tanks B0 and BF respectively which are rigidly connected to each other and to the airframe 32. The boost oxidizer and boost fuel tanks are automatically jettisoned upon conclusion of boost operation phase (FIG. 2) by shaped charges set off by a control signal, as will be described. The rocket engine and its controls designated as a whole as .I is mounted centrally in the aft end of the airframe as is conventional.

Throughout the specification the term engine refers to the combustion chamber including the exhaust nozzle, propellant injection means, etc., while the term powerplant refers to the missile as a whole including the engine, the tankage, conduits, controls, etc.

BASIC FORM OF ACORN ROCKET ENGINE A basic form of the novel rocket engine of the present invention is disclosed in FIG. 6 in which numerals 33 and 34 designate pressurized oxidizer and fuel tanks of the powerplant respectively, the rear wall 35 of the latter being contoured in shape, and in this particular embodiment, substan tially hemispherical. Oxidizer and fuel are delivered to the conical or acom thrust chamber 36 by concentric manifolds 37, 38 on which it is rigidly mounted. The rim of its large open end 40 is spaced from but adjacent the concave surface of the aft wall 35 of the fuel tank 34 so as to define an annular exhaust passage or nozzle throat 41 therewith.

The fuel in passing to the injection orifices 42 of the combustion and thrust chamber is directed along the rear hemispherical wall 35 by a confon'ning bafile 43 so as to regeneratively cool this exhaust gas deflecting surface. Similarly, oxidizer and fuel regeneratively cool the acorn thrust chamber 36 by means of a compartmented coolant jacket 44 before separately reaching the injection orifices 42. The fuel and oxidizer, being hypergolic, require no ignition and mix and burn in the thrust chamber, the gases passing around the open end 40 through the annular nozzle throat 41 and back along the deflecting wall 35, expanding and providing thrust.

It is to be noted that the arrangement just described enables the volume of the missile normally occupied by the combustion chamber to be available for propellant loading, thus effecting an increase in mass ratio for the missile. The use of the aft tank head or wall 35 for the deflection portion of the nozzle also increases the structural efficiency of the airframe so that the tank head serves a dual purpose.

BASIC ACORN ENGINE WITH MECHANICAL TI-IRUST VECTOR CONTROL The basic acorn engine .I of FIG. 6 is readily modified for thrust vector control with mechanical actuating means as shown schematically and in section in FIGS. 7-9 inclusive. The aft fuel tank wall 35 is modified to receive a ring 45 whose aft face is flush, i.e., providing uniform peripheral distribution of throat area, with the surface of the deflecting wall 35 except during those periods when a thrust vector direction other than axial of the motor is desired.

The ring 45 comprises three equal, arcuate sectors 46 to each of which one or more fluid pressure, extensible cylinders 47 are pivotally connected by V-shaped legs 48 at one end and to the central propellant conduit 49 at their other ends. The sectors 46 move in guides 50 to prevent their cocking which is also prevented by the V-shaped legs 48.

When thrust vector control is desired, fluid pressure is introduced into one or more of the cylinders 47 to move one or more of the sectors 46 into the throat area 41 to vary it and cause local disturbance to effect the vector control. Similar control is effected by the means disclosed in FIG. 9 which differs from FIGS. 7 and 8 only in that a single complete ring 53 is used instead of the sectors 46.

It will be appreciated that the arrangements of FIGS. 7-9 inclusive embodies a number of advantages in thrust vector control in that pivoting of the acorn is not required so that there is no resultant axial load which must be coped with, nor need for pivotal propellant lines. Moreover, the central conduit 49 permits efficient hydraulic actuation design.

As indicated above, the requirement for thrust vector control normally requires that the combustion chamber of a rocket engine be swiveled or girnballed, or, as illustrated, that auxiliary devices be used to effect jet deflection. The acorn thrust chamber, as shown in FIG. 10, embodies another new and unique mechanical way of obtaining thrust vector control by providing the concentric fuel and oxidizer manifolds 37, 38 which support the acorn 36 with a ball joint 54 which permits swiveling of the acorn thrust or combustion chamber at, or near, its center of gravity as indic'ated in the dotted lines. A suitable joint seal and a flexible diaphragm 55 prevent leakage or premature mixing of the propellants which pass through the ball joint 54 and enter the cooling jacket 44 as earlier described.

The mounting of the acorn 36 at or near its center of gravity enables the minimizing of the effect of inertial loads during its maneuvering to also enable a minimization of actuating force. As shown in dotted lines in FIG. 10, the acorn 36 is deflected by one of three fluid actuators 56 spaced apart so that the proportion of thrust generated in the upper portion is appreciably greater than that in the lower portion. A clockwise turning (Yawing) movement is therefore generated due to the asymmetric thrust generation.

BASIC ACORN ENGINE WITH THRUST CONTROL As illustrated in FIG. 11, the thrust of the engine .I may be readily varied by mounting the acorn 36 on extensible, concentric manifolds 57 which slide in suitably sealed housing 58 and are connected to a piston 59 in a fluid pressure actuating cylinder 60. Actuation, as demanded by a control system, will linearly translate the acorn chamber 36 along the fore and aft axis to vary the throat area 41 and throttle it as shown at 63 and thus change the thrust of the engine. As the throat area 41 is decreased the expansion area ratio increases which is desirable for high performance of the sustainer at extreme altitudes. 

1. A rocket engine comprising, in combination, a manifold for introducing propellants into a combustion chamber, a wall having a hemispherical deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound, a conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, apertures formed in the apex of said chamber for bleeding combustion gases therethrough, and means for deflecting a portion of said gases to vary the thrust vector of said engine in operation.
 2. A rocket engine as recited in claim 1 wherein said deflecting means comprises a source of gas pressure, plural conduits connecting said source and said apertures, and a valve selectively connecting one of said conduits with a portion of said apertures to deflect a portion of said gases.
 3. A rocket engine as recited in claim 2 wherein said combustion chamber is pivotally mounted at substantially its center of gravity on said manifold and the deflection of a portion of said gases effects a pivoting of said chamber to vary the thrust vector of said engine.
 4. A rocket engine as recited in claim 3 wherein the pivot of said chamber comprises a ball joint in said manifold aft of said surface.
 5. A rocket engine as recited in claim 4, and means for delivering a fluid under high pressure to and between the mating parts of said ball joint to balance the pressures therebetween to permit frictionless pivoting of said chamber.
 6. A rocket engine comprising a manifold for introducing propellants into a combustion chamber, a wall having a hemispherical deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound, a substantially conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, and mechanical means connected with said manifold for varying the thrust vector of said engine in operation, said means comprising a circular slot formed in said wall surface adjacent and concentric with said manifold, a plurality of arcuate abutting segments slidably mounted in said slot flush with said surface, and actuator means connecting said manifold with each of said segments and separately operable to move a segment into the nozzle throat.
 7. A rocket engine comprising a manifold for introducing propellants into a combustion chamber, a wall having a hemispherical deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound, a substantially conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, and mechanical means connected with said manifold for varying the thrust vector of Said engine in operation, said means comprising a circular slot formed in said wall surface adjacent and concentric with said manifold, a ring tiltably and slidably mounted in said slot flush with said surface, and a plurality of actuators connecting said manifold with circumferentially spaced points of said ring and separably operable to move said ring at a selected point into the nozzle throat.
 8. A rocket engine comprising a manifold having an igniter therein, a hemispherical deflecting surface mounted adjacent the forward end of said manifold and extending aft therearound, a conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion nozzle and throat therewith, a solid propellant molded in said chamber and ignitable by said igniter, and mechanical means connected with said manifold for varying the thrust vector of said engine in operation.
 9. A rocket engine as recited in claim 8 wherein said means comprises a circular slot formed in said wall surface adjacent and concentric with said manifold, a plurality of arcuate abutting segments slidably mounted in said slot flush with said surface, and actuator means connecting said manifold with each of said segments and separately operable to move a segment into the nozzle throat.
 10. A rocket engine as recited in claim 8 wherein said means comprises a circular slot formed in said wall surface adjacent and concentric with said manifold, a ring tiltably and slidably mounted in said slot flush with said surface and a plurality of actuators connecting said manifold with circumferentially spaced points of said ring and separably operable to move said ring at a selected point into the nozzle throat.
 11. A rocket engine comprising a manifold having an igniter therein, a hemispherical deflecting surface mounted adjacent the forward end of said manifold and extending aft therearound, a conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion nozzle and throat therewith, and a solid propellant molded in said chamber and ignitable by said igniter, said manifold being axially movable with respect to said surface to vary the throat area of said nozzle, and said manifold including means for permitting pivoting of said chamber at substantially its center of gravity to vary the thrust vector thereof.
 12. A rocket engine comprising a manifold having an igniter therein, a hemispherical deflecting surface mounted adjacent the forward end of said manifold and extending aft therearound, a conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion nozzle and throat therewith, and a solid propellant molded in said chamber and ignitable by said igniter, said manifold being axially movable with respect to said surface to vary the throat area of said nozzle.
 13. A rocket engine as recited in claim 12, and actuator means connected with said manifold to effect axial movement thereof.
 14. A rocket engine comprising a manifold for introducing a liquid propellant into a combustion chamber for combustion therein with a solid propellant, a wall having a hemispherical deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound, a conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, a solid propellant molded in said chamber for combustion with said liquid propellant, and mechanical means connected with said manifold for varying the thrust vector of said engine in operation, said means comprising a circular slot formed in said wall Surface adjacent and concentric with said manifold, a plurality of arcuate abutting segments slidably mounted in said slot flush with said surface, and actuator means connecting said manifold with each of said segments and separately operable to move a segment into the nozzle throat.
 15. A rocket engine comprising a manifold for introducing a liquid propellant into a combustion chamber for combustion therein with a solid propellant, a wall having a hemispherical deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound, a conical combustion chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, a solid propellant molded in said chamber for combustion with said liquid propellant, and mechanical means connected with said manifold for varying the thrust vector of said engine in operation, said means comprising a circular slot formed in said wall surface adjacent and concentric with said manifold, a ring tiltably and slidably mounted in said slot flush with said surface, and a plurality of actuators connecting said manifold with circumferentially spaced points of said ring and separably operable to move said ring at a selected point into the nozzle throat.
 16. A rocket engine comprising a manifold for introducing a monopropellant into a combustion chamber for decomposition therein by a catalyst, a hemispherical wall and deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound, a conical decomposition and thrust chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, a catalyst bed mounted in said chamber for the passage of decomposing monopropellant therethrough, and mechanical means connected with said manifold for varying the thrust vector of said engine in operation, said means comprising a circular slot formed in said wall surface adjacent and concentric with said manifold, a plurality of arcuate abutting segments slidably mounted in said slot flush with said surface, and actuator means connecting said manifold with each of said segments and separately operable to move a segment into the nozzle throat.
 17. A rocket engine comprising a manifold for introducing a monopropellant into a combustion chamber for decomposition therein by a catalyst, a hemispherical wall and deflecting surface fixed to an intermediate portion of said manifold and extending aft therearound a conical decomposition and thrust chamber fixed to and communicating with the aft end of said manifold and having its larger open end positioned adjacent but spaced from said surface to define an expansion exhaust nozzle and throat therewith, a catalyst bed mounted in said chamber for the passage of decomposing monopropellant therethrough, and mechanical means connected with said manifold for varying the thrust vector of said engine in operation, said means comprising a circular slot formed in said wall surface adjacent and concentric with said manifold, a ring tiltably and slidably mounted in said slot flush with said surface, and a plurality of actuators connecting said manifold with circumferentially spaced points of said ring and separably operable to move said ring at a selected point into the nozzle throat. 